Autonomous manoeuring for spinning spacecraft

ABSTRACT

An inventive autonomous active manoeuvring method and system  1  for spinning spacecraft is provided having a capability to enhance the AOCMS performance of passive spinning satellites and to fulfil the emerging autonomy requirements applicable to new generation satellites. In broad terms, the invention resides in (a) the overall concept of providing autonomous execution of spin axis re-orientation manoeuvring for spinning spacecraft designed and executed autonomously on-board the spacecraft by the AOCMS and (b) in the proposed strategy set in place to execute the re-orientation manoeuvres with respect to the handling of residual nutation. Advantageously, the provision of coupling nutation avoidance manoeuvres with active nutation damping  8  on board the spacecraft reduces/minimises the manoeuvre settling time required to return the spacecraft to the steady state pointing performance, while not imposing constraints upon the particular spacecraft inertia sensor properties.

FIELD OF THE INVENTION

The present invention relates to an autonomous active manoeuvring methodand system for spinning spacecraft, and more particularly, but notexclusively, concerns a method and system of autonomous execution ofspin axis re-orientation manoeuvres of spacecraft designed and executedautonomously on-board by AOCMS (that is, estimation guidance andon-board closed loop control).

BACKGROUND OF THE INVENTION

On conventional spinning spacecraft, such as for the Giotto and Clustermissions, the ground is entirely in charge of attitude determination andthe design of momentum re-orientation manoeuvres: the ground performsthe attitude reconstitution through measurements from a (non-autonomous)star mapper and computes the next manoeuvre to be executed in near realtime. The manoeuvre is computed and executed “a priori”, from theknowledge of the actual spin axis attitude, and the prescribed spin axisattitude. This is illustrated in FIG. 1.

This known approach reduces/minimises the on-board AOCMS softwarecomplexity, but at the expense of ground computations, and duringnon-contact periods, manoeuvre realisation errors accumulate. It alsorequires a specific solution to master the nutation generated by themomentum reorientation manoeuvres. In this connection, the followingrequirements are to be noted:

-   (1) No closed loop control is required on-board, allowing for a    simpler AOCMS on-board; closing the loop is performed on-ground. No    on-board autonomous attitude determination is therefore required,    and no on-board closed loop computation and execution of manoeuvres    is therefore required.-   (2) Nutation Avoidance Manoeuvres strategies for momentum    re-orientation manoeuvres are required to maximise the useful    mission time (i.e. when the pointing stabilisation requirement is    met).

In this known strategy, as illustrated in FIG. 2, the manoeuvre is splitinto two small pulses. Nutation is shown to be minimised by phasing thetwo pulses with respect to each other. The thrust angle is a function ofthe spin rate (to perform a thrust in the right direction with respectto the inertial frame), and of the nutation rate (to perform a thrustthat “kills” the nutation of the first thrust in the body frame). Theefficiency of the Nutation Avoidance Manoeuvre is directly given by thethrust angle. Some limitations are however inherent to this knownstrategy. For example, for some spacecraft inertial tensor properties,it is not possible to find both a thrust angle and a thrust phasing thatnull the nutation at the end of the spin axis re-orientation. This kindof open loop Nutation Avoidance Manoeuvre is inefficient for someunfavourable inertia ratios, as shown in FIG. 3, thus, and requires theinertial nutation period to be phased with the body nutation period. Incase no phasing is available (within a short time, typically before twospin periods), there is an unwanted residual nutation at the end of thesecond manoeuvre.

-   (3) Residual nutation control is performed passively, using Passive    Nutation Dampers. Note that a Passive Nutation Damper (PND)    typically comprises a tube and end pots filled with fluid. Nutation    creates cyclic acceleration along the tube and energy dissipation    within the fluid makes the nutation decay, so motion tends to a pure    spin around the principal inertia axis. PNDs are tuned on the    nutation frequency. The tuning is achieved by appropriately    designing their dimensions with respect to the dynamic    characteristics of the spacecraft. The resulting PND time constant    depends on the spacecraft geometry, on the spin rate and on the PND    intrinsic characteristics.

In the framework of the new generation of mission, however, theincreased autonomy requirements (aimed at reducing the operationalcosts) and the more complex spin momentum attitude requirements (aimedat providing sufficient manoeuvrability, as required by the mission skyor Earth coverage) do not make near real time operations using the abovedescribed known approach simple. To fulfil the requirements of the newgeneration missions, attitude determination has to be done on a typical24 hours basis, and manoeuvres have to be computed and executed, “apriori”, from the initial attitude determination solution for,typically, the coming 48 hours.

Further, error accumulation increases as momentum re-orientationmanoeuvre realisation errors accumulate. As an example, when usingthrusters, these are mainly linked to thruster impulse bit repeatabilityand sensitivity to initial actual thruster temperature. Erroraccumulation yields a pointing drift error of typically a few arcminutes per day, depending on the actual correlation between independentmanoeuvres. Usage of a solution, as currently defined for formerspinning satellites, can be detrimental to the required medium termpointing stability for the mission.

From the short term pointing stability point of view, it is desired tominimise the nutation generated by the momentum re-orientationmanoeuvres, and also to minimise the subsequently required time to dampany residual nutation. Both items are needed in order to reach a“steady-state” pointing stability performance as soon as possible. Thisin turn maximises useful mission time. Tightly controlling the nutationprincipally through Nutation Avoidance Manoeuvres using knownarrangements, however, can either cause constraints on the spacecraftmass properties, or require a strategy redesign if the mass propertiesevolve.

On the other hand, the usage of known Passive Nutation Dampers causeaccommodation and validation issues for spinning spacecraft. This isparticularly true when a short duration is available after spacecraftmomentum repainting manoeuvres.

OBJECTS AND SUMMARY OF THE INVENTION

The present invention aims to overcome or at least substantially reducesome of the above-mentioned drawbacks.

It is an object of the present invention to provide an improvedautonomous active manoeuvring method and system for spinning spacecraftwith a capability to enhance the AOCMS performance of passive spinningsatellites, and to fulfil the emerging autonomy requirements, ashereinabove discussed, applicable to new generation satellites.

It is another object of the present invention to alleviate theon-board/on-ground interface for spacecraft related to the execution ofspin axis re-orientation manoeuvres, while reducing/minimising theduration for controlling the nutation generated by these manoeuvres. Thefollowing points are to be noted in this connection:

-   (1) The on-board/on-ground interface defined to achieve the on-board    attitude restitution and manoeuvres stems from both the autonomy and    the pointing requirements applicable to new generation spinning    satellites. A definite improvement is proposed in the present    invention, as compared to previous spinner missions, for the    benefits of operation simplicity and pointing performance.-   (2) Rather than relying exclusively on ground computations for    attitude determination and for manoeuvre definition, autonomous    attitude determination in the present invention is foreseen on board    the spacecraft. Attitude knowledge on board advantageously allows    the AOCMS to autonomously design the amplitude of the manoeuvres,    with the additional objective of correcting execution errors from    former manoeuvres.-   (3) The evolutions in spacecraft definition and requirements, as    well as the availability of new technologies (for instance, laser    gyro), make Active Nutation Damping a preferred solution for control    purposes in the present invention, as compared to the known usage of    Passive Nutation Dampers. This is associated together with Nutation    Avoidance Manoeuvres. Further, the proposed inventive strategy of    the invention minimises/reduces the time required for the spacecraft    to return to the steady state pointing performance, while not    imposing constraints on spacecraft inertia tensor properties.

In broad terms, the present invention resides in (a) the overall conceptof providing autonomous execution of spin axis re-orientation manoeuvresfor spinning spacecraft designed and executed autonomously on-board thespacecraft by the AOCMS (i.e. estimation guidance and on-boardclosed-loop control), and (b) in the proposed strategy set in place toexecute the re-orientation manoeuvres, with respect to the handling ofresidual nutation.

More particularly, the method/system of the present invention takesadvantage of the following techniques, in combination, so as to achievethe desired technical effect, namely:

-   (a) computation of the magnitude and the phase of the initial    thruster pulses such that the manoeuvring can be completed in a time    less than or equal to half the inertial precession period;-   (b) completion of the manoeuvring based on the time from the start    of the manoeuvre and the angular rates on the spacecraft body as    measured using gyroscopes; and-   (c) autonomous attitude determination between manoeuvres such that    an error in a particular manoeuvre (due to thruster misalignments    etc) can be corrected in subsequent manoeuvres.

Therefore, according to the present invention, there is provided anautonomous active manoeuvring method for performing autonomously, inclosed loop on-board, a series of fine attitude manoeuvres for aspinning spacecraft, the method comprising:

-   (a) arranging a plurality of thrusters in a predetermined spatial    configuration at the spacecraft;-   (b) controllably generating a series of thruster pulses by means of    the thrusters in said predetermined spatial configuration, the    magnitude and the phase of the thruster pulses being determined by    computation means so as to permit fine attitude reorientation    manoeuvring of the spacecraft about its spin axis, the manoeuvring    being effected in a time duration less than or equal to half the    inertial precession period of the spacecraft about its spin axis;-   (c) completing the manoeuvring in dependence upon the measured time    from the start of the manoeuvre and the measured angular rates    associated with the spacecraft nutation; and-   (d) providing an autonomous attitude determination of the spacecraft    between successive manoeuvring steps such that, in operation of the    spacecraft, the errors associated with particular manoeuvres can be    controllably corrected in subsequent manoeuvres.

In accordance with an exemplary embodiment of the invention which willbe described hereinafter in detail, active nutation damping (AND) isperformed under AOCMS control by enabling rate controller means on-boardthe spacecraft, the nutation being sensed by a (dedicated) sensor anddamped through (dedicated) control and (dedicated) actuator means. Asalready mentioned, coupling nutation avoidance manoeuvres with activenutation damping reduce/minimise the manoeuvre settling time required toreturn the spacecraft to the steady state pointing performance, whilenot imposing constraints upon the particular spacecraft inertia sensorproperties.

Advantageously, the series of thruster pulses are time-phased in amanner which takes account of variations in the measured angular ratesassociated with the spacecraft nutation, thereby enabling the measuredangular rates to be actively controlled.

Advantageously, the measured angular body rates are measured to highaccuracy by gyroscopic means and the generated thruster torques can beapplied on two predetermined transverse axes of the spacecraft, theseaxes being defined as orthogonal to the spin axis of the spacecraft.

The attitude estimation on-board the spacecraft can be convenientlyperformed through an autonomous star tracker, or alternatively, by usinga non-autonomous attitude sensor, a V-sight star mapper for example.

The present invention further extends to a spacecraft system adapted andarranged to carry out the above described method comprising:

-   a plurality of thrusters arranged in a predetermined spatial    configuration at the spacecraft, the thrusters being operable to    generate controllably a series of thruster pulses;-   means for determining the magnitude and the phase of the thruster    pulses so as to permit fine attitude reorientation manoeuvring of    the spacecraft about its spin axis, the manoeuvring being effected    in a time duration less than or equal to half the inertial    precession period of the spacecraft about its spin axis;-   means for completing the manoeuvring in dependence upon the measured    time from the start of the manoeuvre and the measured angular rates    associated with the spacecraft nutation; and-   means for providing an autonomous attitude determination of the    spacecraft between successive manoeuvring steps such that, in    operation of the spacecraft, the errors associated with particular    manoeuvres can be controllably corrected in subsequent manoeuvres.

The present invention can conveniently be embodied in software.

The above and further features of the invention are set forth withparticularity in the appended claims and will be described hereinafterwith reference to the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic of a known cluster on-ground/on-board task sharingarrangement for attitude determination/reorientation manoeuvres ofspacecraft;

FIG. 2 is a diagram for a conventional open loop nutation avoidancemanoeuvre for spacecraft;

FIG. 3 is a table showing typical inertia ratios for a conventional openloop nutation avoidance manoeuvre for spacecraft;

FIG. 4 is a schematic of an arrangement embodying the present invention;

FIG. 5 is a schematic of an arrangement incorporating a star trackerembodying the present invention;

FIG. 6 is a schematic of an arrangement incorporating a star mapperembodying the present invention;

FIG. 7 is a diagram for the inventive closed loop nutation avoidancemanoeuvre of the present invention;

FIG. 8 shows example simulations for closed loop manoeuvre(s) of thepresent invention;

FIG. 9 is a diagram showing how the spacecraft nutation is cancelledusing thruster actuation in the present invention;

FIG. 10 shows a typical star mapper scan path using the arrangement ofFIG. 6;

FIG. 11 is a flow chart of the steps carried out by the arrangement ofFIG. 4 to determine the spacecraft attitude on-board;

FIG. 12 shows a typical star mapper output using the arrangement of FIG.6;

FIGS. 13 and 14 show example Monte Carlo results for use in theinvention;

FIGS. 15 and 16 show typical sequences of manoeuvres for use in theinvention; and

FIG. 17 shows the drift of the spacecraft spin axis typically arisingfrom manoeuvring prior to its correction in the invention.

DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENTS

In this specification, the term “autonomous” is taken to mean withoutground intervention other than initialisation from ground. Also, theterm “attitude” is used throughout to mean/cover any orientation withrespect to a predetermined reference frame. Also, the terms “magnitude”and “phase” as used in relation to thruster pulses should be taken tomean duration of the pulses and the time when the pulses are to befired.

Referring first to FIG. 4, there is schematically shown therein apreferred spacecraft system 1 adapted and arranged to carry out theclosed loop nutation avoidance manoeuvre (NAM) of the present invention.More particularly, as shown, the spacecraft system 1 comprises anon-board segment 2 which includes a plurality of thrusters 3 arranged ina predetermined spatial configuration (not shown) at the spacecraft, astar mapper component 4, means 5 for performing attitude determinationon-board in closed loop, scanning law algorithm means 6, computationmeans 7 for computing the NAM manoeuvres and executing the manoeuvresclosed loop to target the required spacecraft angular momentumorientation in inertial space and active nutation damping means 8 forproviding closed loop nutation control at the end of the (NAM)manoeuvres. As shown in the Figure, the on-board segment 2 is furtherarranged to co-operate with a remote ground segment 10 by telemetry. Theground segment 10 comprises means for attitude monitoring 11 and meansfor providing attitude information 12, for example momentum attitudeinformation requirements. The on-board segment 2 is coupled to thedynamics part 15 of the spacecraft.

In operation of the spacecraft system 1, therefore, the star mapper 4transmits raw data to ground 10 whereupon an attitude determinationon-ground is made using the ground components 11 and 12. Thereafter, thedetermined attitude information from ground is transmitted to theon-board segment 2 and this transmitted information is used toinitialise a fine attitude estimation on-board and to execute thescanning law algorithm which, in turn, are used to compute the thrusteron-times, stop-times and other relevant data for every manoeuvre to beexecuted. This data is used, in turn, to enable the active nutationdamping means 8, permitting closed loop nutation control to beautomatically applied at the end of each manoeuvre, if desired. Thedescribed operation can be controllably repeated by repeating the stepsin closed loop on-board, namely by operating together the attitudedetermination means 5 and the algorithm means 6 in closed loop, enablingthe errors in prior manoeuvres to be corrected.

Referring to FIG. 5, there is schematically shown therein anotherpreferred spacecraft system 20 adapted and arranged to carry out theclosed loop nutation avoidance manoeuvre (NAM) of the present invention.As shown in the Figure, the spacecraft system 20 comprises an on-boardsegment 21 which includes a plurality of thrusters 23 arranged in apredetermined spatial configuration (not shown) at the spacecraft, astar tracker component 24, means 25 for formatting data and performing afine attitude determination on-board in closed loop, scanning lawalgorithm means 26 and computation means 27 for computing the NAMmanoeuvres and executing the manoeuvres closed loop to target therequired spacecraft angular momentum orientation in inertial space. Asshown, the on-board segment 21 is arranged to co-operate with a remoteground segment 28 by telemetry. The ground segment 28 comprises attitudemonitoring equipment 29.

Advantageously, the star tracker component 24 is an autonomous trackerinsofar as the attitude estimation process is completely handled by thetracker, requiring minimal or no interfaces with the AOCMS. In thisembodiment, the star tracker component 24 comprises new generation unitsusing an embedded star catalogue coupled with acquisition and trackingalgorithms, and the tracker is conveniently designed to send attitudedata to ground 28 at any particular requested time.

In operation of the system 20, the on-board star tracker 24 transmitsdata to the formatting means 25 and this transmitted data isappropriately formatted and used to initialise a fine attitudeestimation on-board and to execute the scanning law algorithm 26 whichin, turn, are used to compute the thruster on-times, stop-times andother relevant data for every manoeuvre to be executed. The describedoperation can be controllably repeated by repeating the above describedsteps in closed loop on-board, and conveniently, the attitude dataon-board can be controllably transmitted to ground attitude monitoringmeans 29 at particular requested times, if desired. In this way, theacquisition of knowledge of the current attitude on-board provides themeans to reduce/avoid accumulation of re-orientation manoeuvre errors,any manoeuvre being able to correct for the previous computed manoeuvrerealisation error.

Conveniently, the arrangement of FIG. 5 using autonomous star trackermeans can be envisaged for use in low spinning spacecraft missions(typically, for low spin rates below 1 rpm).

Referring to FIG. 6, there is schematically shown therein anotherspacecraft system 30 incorporating a non-autonomous star mapperembodying the present invention. The arrangement shown in the Figure isalmost identical to the previously described FIG. 4 arrangement and onlythe essential differences between these two Figures will be describedhereinbelow. For sake of clarity, FIG. 6 employs the same referencenumerals as are employed in FIG. 4 for same/like parts of the spacecraftsystem.

As shown in FIG. 6, the on-board attitude determination is defined, atits simplest level, as the autonomous restitution of a fine attitudefrom well predicted attitude information after a re-orientationmanoeuvre. The activity is performed during periods in betweenre-orientation manoeuvres. The on-board attitude knowledge is propagatedthrough star mapper measurement during the re-orientation manoeuvres.This is achieved by using a non-autonomous star mapper 4 such as aV-slit star mapper. Alternatively, this could be achieved throughgyroscope measurement (not shown). Note that in this arrangement, asshown, the absolute attitude determination stands as a ground activity10, but is envisaged to be limited to an initial attitude determination31 with possibly large errors and to regular checking of the quality ofthe relative attitude determination performed on-board (i.e.verification of no drift). Note that no active nutation damping is shownto be provided on-board in this embodiment (unlike in the FIG. 4embodiment).

It is to be appreciated that the simple proposed solution is notlimitative and can be extended to full autonomous attitude determinationon-board.

Conveniently, the arrangement of FIG. 6 using a non-autonomous starmapper 4 can be used for high spin rates (typically up to 20 rpm).Conveniently, the acquisition of knowledge of the current attitudeon-board provides the means to reduce/avoid accumulation ofre-orientation manoeuvre errors, any manoeuvre being able to correct forthe previous computed manoeuvre realisation error.

Re-Orientation Manoeuvres and Nutation Control

The spin axis scan strategy employed in the invention is an improvementover the known Nutation Avoidance Manoeuvres developed for passivespinning satellites. The use of high accuracy gyroscopes, for example,obviates the need to tightly control the inertia ratio throughout themission life, provides the most efficient manoeuvre, and allows themanoeuvre to be completed in a very short time.

The Spin axis scan law is computed relative to the inertial attitude,and is uploaded regularly, upon ground request, to the AOCMS. This scanlaw is a mathematical law describing the required motion of the spinaxis over the required autonomy period. The AOCMS software computes therequired attitude manoeuvres on board according to the prescribedattitude and the actual attitude: amplitude and phase of spin axis tilt.The phase of spin axis tilt is converted into the spacecraft time atwhich to start the manoeuvre, as derived from the spacecraft thrusterconfiguration.

A more generic strategy than known Nutation Avoidance Manoeuvre isproposed in the invention using gyroscopes to complete the manoeuvre. Inan embodiment of the invention using two thrusters, the magnitude of thefirst of two thruster actuations is calculated based on the thrustertorque level. This provides a change of momentum vector that is in thecorrect direction but is half of the magnitude required. The firstopportunity to complete the manoeuvre is half way through the precessioncycle of the spin axis in inertial space. It is possible to complete themanoeuvre at a time from the start of the manoeuvre given by t=n+0.5) Pwhere P is in the inertial nutation period and n is any positive integeror zero. This is shown in FIG. 7.

The manoeuvre is then completed at the predetermined time by enablingrate controller means at the spacecraft, which controller means is incharge of stopping the precession of the spin axis and cancelling thenutation, as illustrated in FIG. 7. The nutation magnitude θ isestimated from the transverse rates measured by the gyroscopes using thefollowing relationship:$\omega_{{pitch}/{yaw}} = \frac{{I_{spin} \cdot {\omega_{spin}.\tan}}\quad\theta}{I_{{pitch}/{yaw}}}$

The rate controller means calculates the required thruster-on times toremove the transverse angular rates measured by the gyroscopes.

The adapted strategy in the invention bears the advantage that theinertia ratio does not need to be confined to specific regions. Themanoeuvre can always be executed and the nutation minimised whatever theinertia ratio. An example of typical simulation results is given in FIG.8.

Note that a small tilt of the manoeuvre can also be advantageouslycommanded to reduce/minimise the manoeuvre duration. Indeed, thisinvolves a larger nutation (dotted circle in FIG. 7), allowing thetransverse rates to reach the threshold more quickly, and to cancel thenutation in a single thruster actuation (dotted lines in FIG. 9). Theprice associated with this faster strategy is a degradation of themanoeuvre efficiency as the cosine of the tilt angle (a).

It is to be appreciated that the accuracy of a re-orientation manoeuvredepends on the knowledge of the inertia ratios and on the thrustererrors, and the transverse rate thresholds are sized from the followingvalues:${\delta\omega}_{y} = {{\frac{T_{y} \cdot \left( {\lambda_{y} - 1} \right) \cdot T_{on}}{\left( {I_{xx} - I_{yy}} \right)}\quad{\delta\omega}_{z}} = \frac{T_{z} \cdot \left( {\lambda_{y} - 1} \right) \cdot T_{on}}{\left( {I_{xx} - I_{zz}} \right)}}$where T_(y), T_(z) are the thruster pitch and yaw torques,

-   -   γ_(y), γ_(z) are the inertia ratios, and    -   Ton is the minimum on-time of the thrusters.        Technical Result of the Invention        Attitude Estimation

As covering the widest range of spin rate considering existing attitudesensors, the concept of an autonomous AOCMS attitude estimation on aspinning satellite has been prototyped with a non-autonomous starmapper. This attitude estimation solution is only illustrative, with theobjective to demonstrate feasibility, even using non-autonomous sensingunits. This solution appears as a simple candidate solution withexisting hardware, and can be further enhanced though hybridising starmapper attitude data with, for example, three-axis gyrometer measurementdata. It can also be traded-off with the direct usage of an autonomousstar tracker, for low spin rate missions.

In the FIGS. 4 and 6 embodiments of the invention, as shown andpreviously described, a star mapper such as the one used for Giotto andCluster missions is used. This unit is preferably of the “V” slit type,with its optic axis orthogonal or canted with respect to the spin axis.As the spacecraft rotates, the field of view of the sensor scans anannular region of the sky, bounded (in the simplest case) by two circlescentred on the spin axis. This is shown on FIG. 10 (the X-axis isconsidered as the spin axis as a convention). The star mapper outputs aseries of event timings for stars crossing each of the slits. Stars nearthe detection threshold may be registered unpredictably, and some falseevents may contaminate the raw data. It is the task of the on boardsoftware to process the raw event timings to generate an accurateestimate of the spin axis attitude. This level of autonomyadvantageously replaces an activity traditionally done by ground. As aby-product, a highly accurate estimate for the spin rate is generatedfrom the star mapper data.

In the FIGS. 4 and 6 embodiments of the invention, the attitudedetermination technique relies on ground telecommand data for an initialattitude estimate, but then self-propagates without any furtherintervention. The attitude estimated on-ground is performed in a wayidentical to that for Cluster, and the resulting quaternion up-linked tothe spacecraft. Periodic re-initialisations are performed to counter anylong-term drift effects. This level of autonomy avoids the need foron-board star pattern recognition. The initialisation specifies the fullattitude state (angles and rates) at a certain point in the spin motion.An initial spin rate estimate is also required, either from gyro data orground command. The algorithm uses the on-board catalogue to predict theevent timings for both star mapper slits, using the attitude estimateand a “priori” knowledge of the spacecraft ‘wobble’ angles (defining thealignment of the main inertia axes in spacecraft reference frame). Timewindows are defined around the predicted event times. Filtering isapplied to remove events not sufficiently separated in time to safelyprevent false star identification (the filtering result isattitude-dependent, and so the catalogue cannot be pre-filtered for thiscriterion). The widths of the applied time windows are chosen such thatthe uncertainty on wobble angles knowledge, or a residual (small)nutation does not prevent from correct star identification. Theremaining well-separated events define an unambiguous association of asubset of star mapper event times with known stars. Mapper times whichlie far from the predicted slit crossing times are rejected, as beingdue to false events or genuine but non-stellar objects (e.g. brightplanets). Mapper events are retained if the time windows for both slitscontain only one event time. Genuine events for faint stars whichregister in only one slit are thus rejected. A flowchart for thealgorithm, as applied to the on-board segment of FIG. 4 for fineattitude computation, is shown in FIG. 11. Typical expected and measuredevent times over 6 spin periods are further shown in FIG. 12. Moreparticularly, as shown in the Figure, the star mapper outputs a seriesof event pulses for each slit. Good stars are stars with well-separatedevent times which avoid possible misidentifications. This filteringmeans that some of the observed mapper times are rejected, those presentin the lower plot but not in the top one. Note that mapper calibrationdata for the delay time bias as a function of star magnitude isadvantageously used to provide time corrections based on catalogued starmagnitude.

For each predicted event where both slits register the star the exacttimings and spin rate are used to construct a star vector in a principalaxis based non-rotating frame. Multiple stars are used to generate anattitude quaternion estimate using the so-called “q method”.

The on-board star catalogue is conveniently built using standardtechniques taking into account the characteristics of the consideredsensing unit. As an example, for the Giotto/Cluster star mapper unit,due to the low sensitivity of the star mapper, only the brightest stars(estimated to 600) are required to give complete coverage to thedetection limit.

Monte Carlo simulation results using a prototype algorithm for use inthe invention are also shown by way of example in FIGS. 13 and 14. Thisincludes nutation and a simple delay time model, but excludes severaleffects such as disturbance torques and natural nutation damping. Thesolution error is typically less than 0.3′ in each axis at 1σ (onestandard deviation). Note, in this connection, that FIG. 13 shows thetypical results obtained from solving for the attitude in 4000 caseswith random uniformly distributed spin axes, using 2 revolutions ofdata. The same initial body rates are used for each case, with anutation angle of 30″. The null error is due to the (uncompensated)delay time, whilst the nutation angle maps onto pitch. Note also thatthe FIG. 14 plot shows the number of stars matched to the on-boardcatalogue. In this case, the minimum number is 4, which occurred inovercrowded regions of the sky, where detectable stars were rejected asbeing too close together for reliable identification. The mean andmedian number of matches here is 13 and a minimum of 2 stars isrequired.

Autonomous Spin Axis Re-Orientation

The new spin re-orientation manoeuvre strategy of the inventionpresented hereinabove has been applied to the realisation of anautonomous sequence of scan manoeuvres illustrated in FIG. 15. Theprinciple is to send to the spacecraft the guidance profile, i.e. thespecification of the successive orientations of the spin axis relativeto an inertial frame over the next autonomy period. The re-orientationmanoeuvres are then performed autonomously to minimise the spin axispointing error accumulation.

As shown in FIG. 15, the scan sequence to be realised over the followingautonomy period is specified as a timeline of reorientation manoeuvresand the requested change in spin axis orientation relative to thereference inertial frame. The exact manoeuvre and rate control enablingtimes are then autonomously computed on-board according to the residualnutation resulting from the previous manoeuvres.

Note that the time of the pulse is computed according to the currentattitude, which relates to a reference time, and the required directionof motion of the spin axis, which equates to a delta time. To completethe manoeuvre, a second delta time is computed which is n+0.5 times theinertial nutation period. The value of n is either 1 or 0, depending onthe inertia properties and the required efficiency of the manoeuvre. Atthis time after the initial impulse, the gyros are used to calculate thethruster actuations required to null the nutation rates on both the Yand Z-axes (X-axis is here defined as the spin axis).

FIG. 16 shows a sequence of five spin axis re-orientation manoeuvres inan inertial reference frame to which the present invention can beapplied. The plot shows the path of the spacecraft X-axis (spin axis)projected into the inertial YZ plane, which is coincident with thespacecraft YZ plane at zero time. The time between the initial and finalthruster actuation for each manoeuvre is half of the inertial precessionperiod (the spin axis follows an half ellipse), i.e. between 20 and 30seconds in the studied case. The residual nutation at the end of thesequence is in the order of 0.1 arcminutes, as is the error in themanoeuvre itself. The dispersion on the final orientation of the spin isdue to the random walk error resulting from gyro measurement errors, thethrusters errors, and uncertainty on the centre of mass. In thisconnection, it is to be noted that the accuracy of the scan manoeuvre istypically better than 0.4′ and that the accumulation of error iseffectively reduced/eliminated by a forward correction strategy. Thenutation is managed/regulated as part of the manoeuvre strategy,ensuring that the pointing stability of the spacecraft is achieved in aminimum time after the beginning of the manoeuvre.

An error on the inertia ratios knowledge results in errors on thephasing between the first manoeuvre and the rate control application. Infact, the inertial nutation period will not be correctly estimated, andthe rate control will be commanded either too earlier or too later thanrequired. This error, if not corrected, creates an apparent drift of thespin axis in the inertial frame:

The error on the knowledge of the inertia ratio results in an errorafter one manoeuvre of:$\delta = \frac{{\delta_{manoeuvre}.\left( {\pi - {2\alpha}} \right)} \cdot {\Delta\lambda}}{2.\cos\quad\alpha}$where:

-   -   δ_(manoeuvre) is the amplitude of the manoeuvre,    -   α is the tilt of the manoeuvre relative to a nominal direction,        and    -   Δλ is the error on the inertia ratio.

This drift of the spin axis is typically illustrated in FIG. 17 and aspreviously described the attitude information which is derived in theinvention provides effective means for compensating for such manoeuvrerealisation errors using a forward correction strategy.

Having thus described the present invention by reference to severalpreferred embodiments, it is to be appreciated that the embodiments arein all respects exemplary and that modifications and variations arepossible without departure from the spirit and scope of the invention.For example, whilst in the described embodiments two or four thrustersare preferably deployed, the accuracy and efficiency of the manoeuvresin the embodiments could possibly be improved, if desired, by provisionof additional thrusters at the spacecraft. Further, it is to beappreciated that the fine attitude reorientation manoeuvring of thespacecraft could even be effected in a time duration marginally morethan half the inertial precession period of the spacecraft about itsspin axis if desired, but that this would be at the cost of notcompleting the spacecraft manoeuvre in the shortest possible time.Further, whilst in the described embodiments the fine attitudedetermination is performed on-board the spacecraft using a autonomousstar tracker or a non-autonomous star mapper, it is equally possible toinitialise and/or to perform the fine attitude determination usingalternative sensors on-board the spacecraft, for example by using a sunsensor and suitable ephemerous on-board the spacecraft.

It is to be appreciated that the detailed definition of the algorithmmeans used on-board for attitude determination in the describedembodiments is dependent on the actual choice of attitude sensor to beused which will be mission dependent.

It is also to be appreciated that the present invention finds utility invarious space missions including low spinning missions (typically below1 rpm) and higher spin rate missions (typically up to 20 rpm) forspacecraft.

1. An autonomous active manoeuvring method for performing autonomously,in closed loop on-board, a series of fine attitude manoeuvres for aspinning spacecraft, the method comprising: (a) arranging a plurality ofthrusters in a predetermined spatial configuration at the spacecraft;(b) controllably generating a series of thruster pulses by means of thethrusters in said predetermined spatial configuration, the magnitude andthe phase of the thruster pulses being determined by computation meansso as to permit fine attitude reorientation manoeuvring of thespacecraft about its spin axis, the manoeuvring being effected in a timeduration less than or equal to half the inertial precession period ofthe spacecraft about its spin axis; (c) completing the manoeuvring independence upon the measured time from the start of the manoeuvre andthe measured angular rates associated with the spacecraft nutation; and(d) providing an autonomous attitude determination of the spacecraftbetween successive manoeuvring steps such that, in operation of thespacecraft, the errors associated with particular manoeuvres can becontrollably corrected in subsequent manoeuvres.
 2. A method as claimedin claim 1, further comprising enabling rate controller means on-boardthe spacecraft during the manoeuvring at a predetermined time from thestart of the manoeuvre such as to actively control the measured angularrates associated with said spacecraft nutation.
 3. A method as claimedin claim 1, wherein said series of thruster pulses are time-phased in amanner which takes account of variations in the measured angular ratesassociated with said spacecraft nutation, enabling said measured angularrates to be actively controlled.
 4. A method as claimed in claim 1wherein the angular rates are measured by gyroscopic means and thethruster torques are applied on two predetermined transverse axes of thespacecraft, the transverse axes being defined as orthogonal to the spinaxis of the spacecraft.
 5. A method as claimed in claim 1 furthercomprising enabling autonomous star tracker means on-board thespacecraft, which tracker means is arranged to co-operate with thecomputation means, providing an autonomous attitude determination of thespacecraft between manoeuvres and establishing the error of priormanoeuvres.
 6. A method as claimed in claim 1 further comprisingenabling a non-autonomous star mapper means on-board the spacecraft,which mapper means is arranged to co-operate with the computation meansproviding an autonomous attitude determination of the spacecraft betweenmanoeuvres.
 7. A method as claimed in claim 6, further comprisingproviding a coarse attitude determination on-ground and performing afine attitude determination on-board the spacecraft based upon acomparison of the attitude information on-board and the attitudeinformation from ground.
 8. A method as claimed in claim 7, wherein thefine attitude determination on-board is effected upon being initialisedby an attitude determination on-ground.
 9. (canceled)
 10. A spacecraftsystem adapted and arranged to carry out a method as claimed in claim 1.11. A system as claimed in claim 10 comprising: a plurality of thrustersarranged in a predetermined spatial configuration at the spacecraft, thethrusters being operable to generate controllably a series of thrusterpulses; means for determining the magnitude and the phase of thethruster pulses so as to permit fine attitude reorientation manoeuvringof the spacecraft about its spin axis, the manoeuvring being effected ina time duration less than or equal to half the inertial precessionperiod of the spacecraft about its spin axis; means for completing themanoeuvring in dependence upon the measured time from the start of themanoeuvre and the measured angular rates associated with the spacecraftnutation; and means for providing an autonomous attitude determinationof the spacecraft between successive manoeuvring steps such that, inoperation of the spacecraft, the errors associated with particularmanoeuvres can be controllably corrected in subsequent manoeuvres.
 12. Acomputer program which when loaded into a computer will enable it tooperate in a system as claimed in claim
 11. 13. (canceled)